This invention relates to the fabrication of composite articles, and, more particularly, to the fabrication of such articles using a fibrous ceramic reinforcing material that is infiltrated with a ceramic matrix material.
In an aircraft gas turbine (jet) engine, air is drawn into the front of the engine, compressed by a shaft-mounted compressor, and mixed with fuel. The mixture is combusted, and the resulting hot exhaust gases are passed through a turbine mounted on the same shaft. The flow of gas turns the turbine, which turns the shaft and provides power to the compressor. The hot exhaust gases flow from the back of the engine, driving it and the aircraft forwardly.
The hotter the exhaust gases, the more efficient is the operation of the jet engine. There is thus an incentive to raise the exhaust gas temperature. However, the maximum temperature of the exhaust gases is normally limited by the materials used to fabricate the turbine vanes and turbine blades of the turbine. In current engines, the turbine vanes and blades are made of nickel-based superalloys and can operate at temperatures of up to 1900-2100.degree. F.
Many approaches have been used to increase the operating temperature limit of the turbine blades and vanes. The compositions and processing of the materials themselves have been improved. Physical cooling techniques are used. In one widely used approach, internal cooling channels are provided within the components, and cool air is forced through the channels during engine operation.
In another approach, ceramic or ceramic composite materials have been used to fabricate some of the hot section components. Most ceramics have very limited fracture toughness, and therefore ceramic composite materials have been considered for such structures. A ceramic composite material of current interest is silicon carbide fibers embedded in a silicon carbide matrix. In one approach, articles are fabricated by collating silicon carbide fibers on a tool, rigidizing the silicon carbide fibers to form a coated preform, and then producing a silicon carbide matrix in the coated preform by chemical vapor deposition (i.e., chemical vapor infiltration) or melt infiltration.
While operable for the fabrication of many articles, the present inventors have recognized that the current manufacturing process has shortcomings when used in the fabrication of other articles. For example, if the article is quite thick, the production of the matrix is slow or may not be possible. Some hollow articles and articles containing cooling channels, such as turbine blades or vanes, cannot be readily prepared by the conventional procedure.
There is a need for an improved approach to the fabrication of composite articles to allow greater flexibility in the preparation of complex and thick sections. The present invention fulfills this need, and further provides related advantages.